FIG. 1 shows a gas turbine engine 10 comprising an air intake 12 and a propulsive fan 14 that generates two airflows A and B. The gas turbine engine 10 comprises, in axial flow A, an intermediate pressure compressor 16, a high pressure compressor 18, a combustor 20, a high pressure turbine 22, an intermediate pressure turbine 24, a low pressure turbine 26 and an exhaust nozzle. Each turbine 22, 24, 26 comprises rotating turbine rotors 27 and stationary nozzle guide vanes (NGVs) 29. A nacelle 30 surrounds the gas turbine engine 10 and defines, in axial flow B, a bypass duct 32.
The air exiting the combustor 20 is generally at a very high temperature, which generally approaches or exceeds the melting point of the materials used in turbine rotors 27 and NGVs 29. Consequently, relatively cool compressor air from the compressors 16, 18 is used to cool components downstream of the combustor such as the turbine rotors 27 and NGVs 29, thereby preventing damage to the components, and increasing their operating life. The compressor air is passed through an interior of the rotors 27 and/or NGVs 29, and out through holes to provide a cooling air film.
In gas turbine engine design, there is a continuing requirement for improved specific fuel consumption. Specific fuel consumption can be improved (i.e. reduced) by increasing the temperature of the combustion products exiting the combustor (known as the turbine entry temperature (TET). Alternatively or in addition, specific fuel consumption can be improved by increasing the pressure ratio provided by the compressors 14, 16, 18.
However, as TET increases, a larger mass flow of cooling air is required in order to maintain the components downstream of the combustor below their maximum temperature. Furthermore, as the compression ratio of the compressed air increases, so does the temperature of the compressor air. In some cases, the compressor air provided by the high pressure compressor 18 can reach temperatures in excess of 700° C. Consequently, the cooling capacity (i.e. the amount of heat that can be removed by the air from a hot fluid at a given temperature) of a given mass of air compressed by the compressors 16, 18 falls as the compression ratio increases, while the requirement for cooling increases as TET increases. Ultimately, a limit is reached whereby providing further cooling air is ineffective at restoring component operating life, and neither compression ratio nor TET can be increased. Furthermore, air used in cooling is less able to take part in the thermodynamic cycle of the engine. Consequently, excessive use of compressor air for cooling may result in an increase in specific fuel consumption at high TET or compression ratios.
One way to overcome this problem is to cool compressor air used for cooling by passing some or all of the cooling air through a heat exchanger such that the cooling air is in heat exchange relationship with a secondary heat exchange medium comprising a relatively cooler fluid. In a gas turbine engine for an aircraft, suitable secondary heat exchange mediums may comprise air from the bypass duct 32, or fuel used to power the gas turbine engine, such as liquid hydrocarbon based fuel.
One example of such an arrangement is described in EP 0469825 in which bypass air is used as the secondary heat exchange medium. However, repeated sudden exposure of the heat exchanger to large thermal gradients, such as will occur when either cooling air or secondary heat exchange medium is bypassed around the heat exchanger, can induce high thermal stresses in the heat exchanger. This may cause sudden or eventual failure of the heat exchanger after a limited number of cycles. Consequently, there is a requirement to increase the longevity of the heat exchanger in such arrangements.
The present invention seeks to address some or all of the above problems.